USE OF BIASED FABRIC TO IMPROVE PROPERTIES OF SiC/SiC CERAMIC COMPOSITES FOR TURBINE ENGINE COMPONENTS

ABSTRACT

The present invention is a ceramic matrix composite turbine engine component, wherein the component has a direction of maximum tensile stress during normal engine operation. The component comprises a plurality of biased ceramic plies, wherein each biased ply comprises ceramic fiber tows, the tows being woven in a first warp direction and a second weft direction, the second weft direction lying at a preselected angular orientation with respect to the first warp direction, wherein a greater number of tows are woven in the first warp direction than in the second weft direction, and wherein a number of tows in the second weft direction allows the biased plies to maintain their structural integrity when handled. The plurality of biased plies are laid up in a preselected arrangement to form the component, and a preselected number of the plurality of biased plies are oriented such that the orientation of the first warp direction of the preselected number of biased plies lie about in the direction of maximum tensile stress during normal engine operation. A coating is applied to the plurality of biased plies. The coating is selected from the group consisting of BN, SiC, and combinations thereof. A ceramic matrix material lies in interstitial regions between the tows of each biased ply and the interstitial region between the biased plies.

This application is a Divisional of U.S. patent application Ser. No.10/784,734, filed Feb. 23, 2004, the contents of which are incorporatedherein by reference.

This invention was made with government support under Contract No.N00421-00-3-0536. The government may have certain rights to theinvention.

FIELD OF THE INVENTION

The present invention relates generally to ceramic matrix turbine enginecomponents, and more particularly, to a ceramic matrix composite turbineblade.

BACKGROUND OF THE INVENTION

In order to increase the efficiency and the performance of gas turbineengines so as to provide increased thrust-to-weight ratios, loweremissions and improved specific fuel consumption, engine turbines aretasked to operate at higher temperatures. As the higher temperaturesreach and surpass the limits of the material comprising the componentsin the hot section of the engine and in particular the turbine sectionof the engine, new materials must be developed.

As the engine operating temperatures have increased, new methods ofcooling the high temperature alloys comprising the combustors and theturbine airfoils have been developed. For example, ceramic thermalbarrier coatings (TBCs) were applied to the surfaces of components inthe stream of the hot effluent gases of combustion to reduce the heattransfer rate and to provide thermal protection to the underlying metaland allow the component to withstand higher temperatures. Theseimprovements helped to reduce the peak temperatures and thermalgradients. Cooling holes were also introduced to provide film cooling toimprove thermal capability or protection. Simultaneously, ceramic matrixcomposites were developed as substitutes for the high temperaturealloys. The ceramic matrix composites (CMCs) in many cases provided animproved temperature and density advantage over the metals, making themthe material of choice when higher operating temperatures were desired.

A number of techniques have been used in the past to manufacture turbineengine components, such as turbine blades using ceramic matrixcomposites. However, such turbine components, under normal operatingconditions, experience varying degrees of local stresses. In thedovetail section of turbine blade components, relatively higher tensilestress regions are located in the outermost portion of the dovetailsection. Ideally, the CMC component would be designed such that thecomponent was stronger in the region of the local stresses. One methodof manufacturing CMC components, set forth in U.S. Pat. Nos. 5,015,540;5,330,854; and 5,336,350; incorporated herein by reference and assignedto the assignee of the present invention, relates to the production ofsilicon carbide matrix composites containing fibrous material that isinfiltrated with molten silicon, herein referred to as the Silcompprocess. The fibers generally have diameters of about 140 micrometers orgreater, which prevents intricate, complex shapes, such as turbine bladecomponents, to be manufactured by the Silcomp process.

Another technique of manufacturing CMC turbine blades is the methodknown as the slurry cast melt infiltration (MI) process. A technicaldescription of such a slurry cast MI method is described in detail inU.S. Pat. No. 6,280,550 B1, which is assigned to the Assignee of thepresent invention and which is incorporated herein by reference. In onemethod of manufacturing using the slurry cast MI method, CMCs areproduced by initially providing plies of balanced two-dimensional (2D)woven cloth comprising silicon carbide (SiC)-containing fibers, havingtwo weave directions at substantially 90° angles to each other, withsubstantially the same number of fibers running in both directions ofthe weave. By “silicon carbide-containing fiber” is meant a fiber havinga composition that includes silicon carbide, and preferably issubstantially silicon carbide. For instance, the fiber may have asilicon carbide core surrounded with carbon, or in the reverse, thefiber may have a carbon core surrounded by or encapsulated with siliconcarbide. These examples are given for demonstration of the term “siliconcarbide-containing fiber” and are not limited to this specificcombination. Other fiber compositions are contemplated, so long as theyinclude silicon carbide.

A major challenge in this approach is fiber coatings that are notuniform, due to the large number of cross-over points in the 2D balancedfabric. The boron nitride (BN) and SiC coatings that are applied to suchbalanced cloth, prior to slurry casting and silicon melt infiltration,are not consistently uniform due to the inability of the coatings toadhere readily to the large number of fiber cross-over points. Inaddition, fiber crimp, which is caused by weaving, causes the loss ofin-plane properties, which diminishes the ability of the CMC componentto endure higher local tensile stresses in the direction of the plane ofthe ply.

In addition, problems with high cycle fatigue (HCF) have been found toresult in CMC component failure when the critical modes of vibration ofcurrent CMC components are within the operating range of the turbineengine. What is needed is a method of manufacturing CMC turbine enginecomponents that permits a more uniform fiber coating. In addition, amethod of manufacturing that addresses the local stress regions and/orensures that the critical modes of vibration of the CMC turbine blade inthe engine environment is not within the operating range of the engineis also needed.

SUMMARY OF THE INVENTION

Improvements in manufacturing technology and materials are the keys toincreased performance and reduced costs for many articles. As anexample, continuing and often interrelated improvements in processes andmaterials have resulted in major increases in the performance ofaircraft gas turbine engines, such as the improvements of the presentinvention. The present invention is a novel method for manufacturing aturbine blade made from a ceramic matrix composite (CMC) using biasedceramic fabric rather than balanced ceramic fabric. The presentinvention produces a component that is stronger in the direction of thehigher tensile stress that is found within discrete higher tensilestress regions within the component during normal engine operation,thereby improving the functionality of the component. In addition, thenovel method of the present invention is useful in ensuring that thecritical modes of vibration of the turbine blade in the turbine engineenvironment are not within the operating range of the turbine engine.

The present invention is a ceramic matrix composite turbine enginecomponent, wherein the component has a direction of maximum tensilestress during normal engine operation. The component comprises aplurality of biased ceramic plies, wherein each biased ply comprisesceramic fiber tows, the tows being woven in a first warp direction and asecond weft direction, the second weft direction lying at a preselectedangular orientation with respect to the first warp direction, wherein agreater number of tows are woven in the first warp direction than in thesecond weft direction, and wherein a number of tows in the second weftdirection allows the biased plies to maintain their structural integritywhen handled. The plurality of biased plies are laid up in a preselectedarrangement to form the component, and a preselected number of theplurality of biased plies are oriented such that the orientation of thefirst warp direction of the preselected number of biased plies lie aboutin the direction of maximum tensile stress during normal engineoperation. A coating is applied to the plurality of biased plies. Thecoating is selected from the group consisting of BN, SiC, andcombinations thereof. A ceramic matrix material lies in interstitialregions between the tows of each biased ply and the interstitial regionbetween the biased plies.

The present invention is also a ceramic matrix composite turbine enginecomponent, wherein the component has a direction of maximum tensilestress during normal engine operation, and comprises a plurality ofceramic plies, wherein each ply comprises ceramic fiber tows, andwherein the tows in each ply lie adjacent to one another in a planararrangement such that each ply has a unidirectional orientation. Acoating is applied to the plies. The coating is selected from the groupconsisting of BN, Si₃N₄, and combinations thereof. The plurality ofplies are laid up in a preselected arrangement to form the component,wherein a preselected number of the plurality of plies are oriented suchthat the orientation of the preselected number of the plurality of plieslie in the direction of maximum tensile stress during normal engineoperation. A ceramic matrix material lies in interstitial regionsbetween the tows of each ply and the interstitial region between theplurality of plies.

The present invention is also a method of manufacturing a turbine enginecomponent, the component having a direction of maximum tensile stressduring normal engine operation, comprising several steps. The first stepis providing a plurality of biased ceramic plies, each biased plycomprising ceramic fiber tows, the tows woven in a first warp directionand a second weft direction, the second weft direction lying at apreselected angular orientation with respect to the first warpdirection, wherein a greater number of tows are woven in the first warpdirection than in the second weft direction, and wherein a number oftows in the second weft direction allows the biased plies to maintaintheir structural integrity when handled. The next step is laying up theplurality of biased plies in a preselected arrangement to form acomponent shape, wherein a preselected number of the plurality of biasedplies are oriented such that the orientation of the first warp directionof a preselected number of the plurality of biased plies lie about inthe direction of maximum tensile stress during normal engine operation.The next step is rigidizing the component shape with a layer of BN and alayer of SiC to form a coated component preform using chemical vaporinfiltration. The next step is partially densifying the coated componentpreform using carbon-containing slurry. The final step is furtherdensifying the coated component preform with at least silicon to form aceramic matrix composite aircraft engine component with biasedarchitecture.

The present invention is a method of manufacturing a CMC turbine enginecomponent with “biased” architecture using the slurry cast MI method. Asdefined herein, “biased” cloth plies have more fiber tows running in afirst direction of the weave of the cloth, the warp direction, than in asecond direction of the weave of the cloth, the weft direction. Thebiased cloth plies should have a ratio of warp fiber tows to weft fibertows of at least about 2:1. Because of this bias, the warp direction ofthe fabric has greater tensile strength in the final CMC product thanthe weft direction. In the method of the present invention a preselectednumber of biased SiC-containing ceramic cloth plies are first laid up toform a turbine engine component shape, such that outer plies, which passthrough regions of higher tensile stress, provide greater tensilestrength. The shape is then rigidized with coatings of BN and SiC usingchemical vapor infiltration (CVI) to form a coated turbine blade preformas known in the art. The preform is then partially densified with acarbon containing slurry as known in the art. The preform is thenfurther densified with silicon to form a CMC turbine engine componentwith biased architecture as known in the art. CMC components that may bemanufactured using such biased architecture include a turbine blade, acooled turbine nozzle, and an uncooled turbine nozzle. In addition toSiC containing ceramic cloth plies, any other type of ceramic clothplies that can be used to form CMC turbine engine components may be usedwith the method of the present invention.

The present invention is also a CMC turbine engine component with biasedarchitecture. Such CMC turbine engine components with biasedarchitecture include a turbine blade, a cooled turbine nozzle, and anuncooled turbine nozzle.

The present invention is also a method of manufacturing a ceramic matrixcomposite aircraft engine component, the component having a direction ofmaximum tensile stress during normal engine operation, comprisingseveral steps. The first step is providing a plurality of prepregceramic plies, the plies comprising prepreg ceramic fiber tows, the towsin each ply lying adjacent to one another in a planar arrangement suchthat each ply has a unidirectional orientation. The next step is layingup the plurality of prepreg ceramic cloth plies in a preselectedarrangement to form a component shape such that a preselected number ofoutermost plies are oriented at about 0° with respect to the directionof maximum tensile stress of the turbine engine component during normalengine operation. The next step is heating the turbine blade shape toform a ceramic preform. The final step is densifying the turbine bladepreform with at least silicon to form a ceramic matrix composite turbineblade.

The present invention is also a method of manufacturing a CMC turbineblade with biased architecture using the “prepreg” MI method. First apreselected number of SiC prepregged plies are laid up in a preselectedarrangement, such that a preselected number of the outermost plies areoriented about at 0°, forming a turbine blade shape. By “0° orientation”with respect to a prepreg ply, it is meant that a ply is laid up suchthat the line of the fiber tows is in the line of the long dimension oraxis of the turbine blade as known in the art. A 90° orientation meansthat the ply is laid up such that line of the fiber tows isperpendicular to the long dimension or axis of the turbine blade asknown in the art. All orientations other than 0° and 90° may be negativeor positive depending on whether the ply is rotated clockwise (positive)from a preselected plane in the long dimension of the turbine blade orrotated counterclockwise (negative) from the preselected plane in thelong dimension of the turbine blade as known in the art. Prepreg pliesthat are oriented at 0° have tensile strength in the final CMC productthat is about twenty times greater than prepreg plies that are orientedat 90°. Such “prepregged” plies comprise silicon-carbide-containingfibers, where the fibers are bundled into tows and the tows are alladjacent to one another such that all of the fibers are oriented in thesame direction. Exemplary processes for making such SiC/SiC prepregmaterial are described in U.S. Pat. Nos. 6,024,898 and 6,258,737, whichare assigned to the Assignee of the present invention and which areincorporated herein by reference. The next step of the process isforming a ceramic preform by heating the turbine blade shape bycompression molding, bladder molding, or autoclaving as known in theart. The final step of the process is densifying the preform withsilicon to form a CMC turbine blade as known in the art.

CMC components that may be manufactured using the prepreg process of thepresent invention include a turbine blade, a cooled turbine nozzle, andan uncooled turbine nozzle. In addition to SiC containing prepregceramic cloth plies, any other type of prepreg ceramic cloth plies thatcan be used to form CMC turbine engine components may be used with themethod of the present invention.

The present invention also provides a CMC turbine engine componentmanufactured with the prepreg process of the present invention. Such CMCturbine engine components with biased architecture include a turbineblade, a cooled turbine nozzle, and an uncooled turbine nozzle.

An advantage of the present invention is that the use of biased fabricin the slurry cast MI process results in significant fiber tow spreadingduring the step of rigidizing, which facilitates more uniform coating ofthe fibers with BN and SiC leading to improved mechanical properties andimproved fracture toughness of the composite.

Another advantage of the present invention is that the use of biasedfabric in the slurry cast MI process results in a lower number of fibercross-over points in the fabric, which results in enhanced in-planemechanical properties.

Another advantage of the present invention is that the use of biasedfabric in the slurry cast MI process allows the tensile strength of theCMC composite to be tailored on a ply-by-ply basis, since the biasedfabric has a greater tensile strength in the warp direction than theweft direction when processed into a CMC component.

Another advantage of the present invention is that the use of specificorientations of prepreg plies on a ply by ply basis allows the tensilestrength of the CMC composite to be tailored, since the prepreg ply hasa greater tensile strength in the 0° orientation than the 90°orientation when processed into a CMC component.

Another advantage of the present invention is that a CMC turbine badecan be manufactured so that the portions of the CMC turbine blade withhigher tensile stresses can have plies in those regions with fibersrunning in the loading direction of the stresses, increasing the crackresistance of the CMC.

Another advantage of the present invention is that the modulus ofelasticity of a CMC component can be tailored by selecting theorientation of the plies on a ply basis, which will ensure that thecritical modes of vibration in an engine environment are not within theoperating range of the of the CMC component.

Other features and advantages of the present invention will be apparentfrom the following more detailed description of the preferredembodiment, taken in conjunction with the accompanying drawings whichillustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an exemplary example of a LPT blade in an aircraft engine.

FIG. 2 is a flow chart illustrating a slurry cast MI method ofmanufacture of the present invention to produce a CMC turbine blade.

FIG. 3 a flow chart illustrating a prepreg MI method of manufacture ofthe present invention to produce a CMC turbine blade.

FIG. 4 is a cross-sectional view of a CMC LPT blade dovetail of thepresent invention manufactured using a prepreg MI method of manufactureand showing the higher tensile stress region and the lower tensilestress region.

FIG. 5 is a cross-sectional view of a CMC LPT blade dovetail of thepresent invention manufactured using a prepreg MI method of manufactureand showing the outermost plies of the dovetail.

FIG. 6 is a cross-sectional view of a CMC LPT blade dovetail of thepresent invention manufactured using a slurry cast MI method ofmanufacture and showing the higher tensile stress region and the lowertensile stress region.

FIG. 7 is a portion of a biased ceramic ply used in the manufacture ofthe slurry cast MI method of CMC component manufacture.

FIG. 8 is a cross-sectional view of a CMC LPT blade dovetail of thepresent invention manufactured using a prepreg MI method of manufactureand showing the outermost plies of the dovetail.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 depicts an exemplary aircraft engine LPT blade 20. In thisillustration a turbine blade 20 comprises a ceramic matrix compositematerial. The turbine blade 20 includes an airfoil 22 against which theflow of hot exhaust gas is directed. The turbine blade 20 is mounted toa turbine disk (not shown) by a dovetail 24 that extends downwardly fromthe airfoil 22 and engages a slot of similar geometry on the turbinedisk. The LPT blade 20 of the present invention does not include anintegral platform. A separate platform is provided to minimize theexposure of the dovetail 24 to hot gases of combustion. The airfoil maybe described as having a root end 40, and an oppositely disposed tip end32.

Referring now to FIG. 2 there is a shown flow chart illustrating aslurry cast MI method of manufacture of the present invention to producea CMC turbine blade. The initial step 100 of the process is laying up apreselected number of biased SiC containing cloth plies of preselectedgeometry in a preselected arrangement to form a turbine blade shape. Ina preferred embodiment, there are a preselected number of fiber towswoven in the weft direction sufficient to allow the SiC cloth to behandled and laid up without falling apart. A CMC element manufacturedwith biased SiC containing cloth plies has greater tensile strength inthe warp direction of the SiC containing cloth plies than the weftdirection. The tensile strength in the warp direction is up to about 25percent greater than in the weft direction.

As the outermost region of a turbine blade dovetail has known regions oflocalized higher tensile stress in the direction of the long dimensionor axis of the turbine blade during engine operation, the fact that thebiased cloth plies have a greater tensile strength in the warp directionthan in the weft direction is used in the selection of the orientationof the outermost plies of the turbine blade shape. A preselected numberof the outermost plies of the turbine blade shape are laid up at anorientation of about 0° so that the plies that pass through regions oflocalized higher tensile stress have more fiber tows that run in theloading direction of the stresses than the opposite direction. Theoutermost plies that lie within the regions of localized higher tensilestress must have more plies laid up at an orientation of about 0° thanlaid up in other orientations. Such orientation increases the fracturetoughness of the final CMC component. By “0° orientation”, it is meantthat the SiC containing cloth is laid up so that the warp direction ofthe cloth is in the line of the axis of the turbine blade as known inthe art. A 90° orientation means that the cloth is laid up so that thewarp direction of the cloth is in the line perpendicular to the longaxis of the turbine blade as known in the art. The All orientationsother than 0° and 90° may be negative or positive depending on whetherthe ply is rotated clockwise (positive) from a preselected plane or axisin the long dimension of the turbine blade or rotated counterclockwise(negative) from the preselected plane in the long dimension of theturbine blade as known in the art. The remaining plies that do not passthrough the regions of localized higher tensile stress may be arrangedin any appropriate orientation as known in the art. For example, theremaining plies could all be laid up in an alternating formation suchthat the remaining plies are at about a 45° orientation, followed byabout a −45° orientation, followed by about a 45° orientation, followedby a −45° orientation, etc. as is known in the art, or such that theremaining plies are at a −45° orientation, followed by a 0° orientation,followed by a +45° orientation, followed by a 90°, or in any othermechanically acceptable arrangement.

In addition, it is well known in the art that the HCF response of aturbine blade is controlled by the modulus of the material used tomanufacture the turbine blade. As the biased fabric allows control overthe directional tensile strength of the CMC turbine blade on a ply byply basis, the weave of the individual plies can be controlled, and theplies may be laid up so that the modulus of elasticity of the densifiedCMC turbine blade is such that the critical modes of vibration in theCMC turbine blade are not within the operating range of the turbineengine. In addition, the lower number of cross-over points in the fabricresult in enhanced in-plane mechanical properties as more of the fibertows are oriented in a more linear direction.

Once the plies are laid up, the next step 120 is rigidizing the turbineblade shape by applying BN and SiC coatings using a chemical vaporinfiltration (CVI) process as is known in the art, forming a rigidcoated turbine blade preform. The use of a larger number of fibersrunning in the warp direction of the cloth reduces the number of fibercross-over points in the fabric, which enables more consistently uniformcoatings of BN and SiC. In addition, the use of biased SiC containingfabric results in fiber tow spreading during the step of rigidizing asthe tows are not held as tightly together as they are in balanced SiCcontaining fabric. Such spreading, however, is not so extensive as tocause the fabric to lose its form. The spreading facilitates moreuniform coating of the fibers with BN and SiC. Such uniform coatingprovides the final CMC component with improved mechanical properties,including improved modulus of elasticity, improved tensile strength, andimproved fracture toughness.

The next step 130 is partially densifying the coated turbine bladepreform by introducing a carbon-containing slurry, as is known in theart, into the porosity of the coated turbine blade preform. The finalstep 140 is further densifying the turbine blade preform with at leastsilicon, and preferably boron doped silicon, through an MI process, asknown in the art, forming a SiC/SiC CMC turbine blade with biasedarchitecture.

In another embodiment of the present invention, the first step 200 ofthe present invention is laying up a preselected number of prepregfabric plies in a preselected arrangement to form the shape of a turbineblade, where a preselected outer number of prepreg plies are laid up inthe 0° orientation. A prepreg ply that has an orientation of 0° withrespect to the tensile stresses has a tensile strength that is abouttwenty times greater than a prepreg ply that has an orientation of 90°with respect to the tensile stresses in the final CMC material. As theoutermost region of turbine blade dovetails have known regions oflocalized higher tensile stress in the direction of the axis of theturbine blade during engine operation, the fact that the 0° prepregplies have a greater tensile strength than the 90° prepreg plies is usedin the selection of the orientation of the outermost plies of theturbine blade shape.

A preselected number of the outermost plies of the turbine blade shapeare laid up at an orientation of 0° to the tensile load so that theplies that pass through the regions of localized higher tensile stresshave fiber tows that run in the loading direction of the stresses. Theoutermost plies that lie within the regions of localized higher tensilestress must have more plies laid up an orientation of about 0° than laidup in other orientations. The remaining plies that do not pass throughthe regions of localized higher tensile stress may be arranged in anyappropriate orientation as known in the art. For example, the remainingplies could all be laid up in an alternating formation such that theremaining plies are at a 45° orientation, followed by −45° orientation,followed by a 45° orientation, followed by a −45° orientation, etc. asknown in the art.

It is well known in the art that the high cycle fatigue (HCF) responseof a turbine blade is controlled by the modulus of the material. As theorientation of the prepreg plies allows control over the stiffness ofthe CMC turbine blade on a ply by ply basis, the orientation of theindividual plies and the plies themselves may be arranged so that themodulus of elasticity of the final CMC turbine blade is such that thecritical modes of vibration in the CMC turbine blade are not within theoperating range of the turbine engine.

The next step 210 is heating the turbine blade shape as known in the artto form a ceramic turbine blade preform by compression molding, bladdermolding, or autoclaving as known in the art. The final step 220 isdensifying the turbine blade preform with at least silicon, andpreferably boron doped silicon, through an MI process, as known in theart, forming a densified SiC/SiC CMC turbine blade. The presentinvention is also a method of manufacturing a turbine engine componentwherein a plurality of plies of biased ceramic cloth are provided andlaid up in a preselected arrangement to form a turbine engine componentshape. The turbine engine component shape is then rigidized to form acoated turbine blade preform. The coated turbine blade preform is thenpartially densified using carbon-containing slurry. The turbine bladepreform is then further densified with at least silicon to form aceramic matrix composite aircraft engine component with biasedarchitecture. The biased ceramic cloth plies may be silicon carbidecontaining plies.

As shown in FIG. 4, which is an example of a cross-sectional view of aCMC LPT blade 20 dovetail 22 of the present invention manufactured withthe prepreg MI process, the two higher tensile stress regions 24 and thelower tensile stress region 26 are evident. A dashed line shows theseparation between the stress regions 24 and 26. The blade 20 comprisesa plurality of ceramic prepreg plies 44 within a ceramic interstitialmatrix 50. FIG. 5, is a cross-section of the outermost plies 44 withinthe higher tensile stress region 24 taken at line 5-5 in FIG. 4. Eachprepreg CMC ply 44 comprises ceramic prepreg tows 46, the coating 48 onthe tows 46, and the interstitial ceramic matrix 50 between the tows 46and plies 44. As can be seen from FIG. 5, more of the plies 44 are 0°oriented plies 52 than are 90° oriented plies 54.

As shown in FIG. 6, which is an example of a cross-sectional view of aCMC LPT blade 30 dovetail 32 of the present invention manufactured withthe slurry cast MI process, the two higher tensile stress regions 34 andthe lower tensile stress region 36 are evident. A dashed line shows theseparation between the stress regions 34 and 36. The blade 30 comprisesa plurality of biased ceramic slurry cast plies 64 within aninterstitial ceramic matrix 68. FIG. 7 shows one example of a biasedceramic cloth ply 58 prior to lay up. As can be seen from FIG. 8, whichis a portion of a biased ceramic cloth ply 58, the biased ceramic ply 58comprises more warp ceramic tows 60 than weft ceramic tows 62, with aratio of at least 2 warp ceramic tows 60 to every weft ceramic tow 62.This is also evident in FIG. 8, which is a cross-section of theoutermost plies 64 within the higher tensile stress region 34 taken atline 8-8 in FIG. 6. Each biased slurry cast CMC ply 64 comprises warpceramic tows 60 and weft ceramic tows 62. The coating 66 on the tows 60,62, and the interstitial ceramic matrix between the tows 60, 62 andplies 64. As can be seen from FIG. 8, more of the plies 64 are 0°oriented plies 64 than are 90° oriented plies 66.

The present invention also includes a ceramic matrix composite turbineengine component, such as a turbine blade, a cooled turbine nozzle, oran uncooled turbine nozzle, wherein the component is initially laid upin a preselected arrangement using a preselected number of biasedceramic fabric plies. The biased ceramic fabric may be silicon carbidecontaining fabric.

The present invention also includes a ceramic matrix composite turbineengine component, such as a turbine blade, a cooled turbine nozzle, oran uncooled turbine nozzle, wherein the component is initially laid upin a preselected arrangement using a preselected number of prepregceramic plies, wherein at least some of the prepreg plies are orientedsuch that the orientation of the fiber tows in the at least some of theprepreg plies is substantially in the line of the direction of thetensile stress of the turbine engine component during normal engineoperation. The prepreg plies may be silicon carbide containing prepregplies.

While the invention has been described with reference to a preferredembodiment, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment disclosed as the best modecontemplated for carrying out this invention, but that the inventionwill include all embodiments falling within the scope of the appendedclaims.

1. A ceramic matrix composite turbine engine component, the componenthaving a direction of maximum tensile stress during normal engineoperation, comprising: a plurality of biased ceramic plies, each biasedply comprising ceramic fiber tows, the tows woven in a first warpdirection and a second weft direction, the second weft direction lyingat a preselected angular orientation with respect to the first warpdirection, wherein a greater number of tows are woven in the first warpdirection than in the second weft direction, and wherein a number oftows in the second weft direction allows the biased plies to maintaintheir structural integrity when handled; the plurality of biased plieslaid up in a preselected arrangement to form the component, wherein apreselected number of the plurality of biased plies are oriented suchthat the orientation of the first warp direction of the preselectednumber of biased plies lie about in the direction of maximum tensilestress during normal engine operation; a coating applied to theplurality of biased plies, the coating selected from the groupconsisting of BN, SiC, and combinations thereof; and a ceramic matrixmaterial lying in interstitial regions between the tows of each biasedply and the interstitial region between the biased plies.
 2. The ceramicmatrix composite turbine engine component of claim 1, wherein theceramic matrix material is silicon carbide.
 3. The ceramic matrixcomposite turbine engine component of claim 1, wherein a ratio of anumber of tows in the first warp direction to the number of tows in thesecond weft direction in each of the biased ceramic plies is at leastabout 2:1.
 4. The turbine engine component of claim 2, wherein thecomponent is a turbine blade.
 5. The turbine engine component of claim2, wherein the component is a cooled turbine nozzle.
 6. The turbineengine component of claim 2, wherein the component is an uncooledturbine nozzle
 7. The ceramic matrix composite turbine engine componentof claim 6, wherein a ratio of a number of tows in the first warpdirection to the number of tows in the second weft direction in each ofthe biased ceramic fiber plies is at least about 2:1.
 8. A ceramicmatrix composite turbine engine component, the component having adirection of maximum tensile stress during normal engine operation,comprising: a plurality of ceramic plies, each ply comprising ceramicfiber tows, the tows in each ply lying adjacent to one another in aplanar arrangement such that each ply has a unidirectional orientation;a coating applied to the plies, the coating selected from the groupconsisting of BN, Si₃N₄, and combinations thereof; the plurality ofplies laid up in a preselected arrangement to form the component,wherein a preselected number of the plurality of plies are oriented suchthat the orientation of the preselected number of the plurality of plieslie in the direction of maximum tensile stress during normal engineoperation; and a ceramic matrix material lying in interstitial regionsbetween the tows of each ply and the interstitial region between theplurality of plies.
 9. The ceramic matrix composite turbine enginecomponent of claim 8, wherein the ceramic matrix material is siliconcarbide.
 10. The turbine engine component of claim 9, wherein thecomponent is a turbine blade.
 11. The turbine engine component of claim9, wherein the component is a cooled turbine nozzle.
 12. The turbineengine component of claim 9, wherein the component is an uncooledturbine nozzle